An end wall cooling structure suitable for high-pressure turbine blades

Abstract

The present invention provides an endwall cooling structure suitable for high-pressure turbine blades, comprising an annular endwall and an endwall mounting edge for mounting the annular endwall. Discrete air film holes are rationally arranged in the area on the inner side of the endwall that is most affected by the mainstream horseshoe vortex, while enhancing the impact heat exchange effect on the outer side of the endwall. This allows for efficient utilization of cold air and provides effective cooling protection for the endwall near the leading edge of the blade. In particular, for endwalls with shorter air inlet edges, which are more difficult to cool, air is introduced through the endwall mounting edge to ensure sufficient cold air flow to the endwall near the leading edge of the blade, thereby enhancing the reliability of endwall cooling.

Description

Technical Field

[0001] The present invention belongs to the technical field of aero-engine turbine blades, and in particular relates to an end wall cooling structure suitable for high-pressure turbine blades. Background Art

[0002] To improve engine performance, it's necessary to increase the compressor's compression ratio and turbine inlet gas temperature. As turbine inlet gas temperature increases, the engine's high-temperature components must withstand greater thermal loads. According to existing literature, the turbine inlet temperature of existing aircraft engines with a thrust-to-weight ratio of 10 reaches 1800K to 1900K. In the future, turbine inlet temperatures for aircraft engines with even higher thrust-to-weight ratios could reach as high as 2100K to 2300K, far exceeding the heat resistance limits of current turbine materials. The endwalls of high-pressure turbine blades face the mainstream high-temperature gas flow. Limited by the confined space of their inherent geometry, complex external flow, and lower cooling airflow compared to the blade body, cooling the turbine endwalls is an increasingly prominent issue.

[0003] According to literature review, most current endwall cooling solutions use a combination of impingement and film holes for endwall cooling. However, as the temperature of aircraft engine combustion gas continues to rise, high-temperature zones are prone to forming on the endwall near the leading edge of the blade. The existing cooling structure still has some problems: 1) The endwall near the leading edge of the blade is not independently partitioned, resulting in poor film coverage from the film holes arranged in this area, and the cooling relies more on the cold air leaking from the slots between the combustion chamber and the turbine; 2) The impingement holes are mostly evenly distributed, and there is no effective impingement heat exchange design for the high-temperature area; 3) For endwall structures with short air inlet edges, due to the limited geometric space, the air inlet position of the cold air is single, and the amount of cold air is insufficient, making the above problems more prominent. Summary of the Invention

[0004] In order to solve the above problems, the purpose of the present invention is to provide an end wall cooling structure suitable for high-pressure turbine blades to solve the problem of insufficient cooling of the end wall near the leading edge of the blade of an aircraft engine.

[0005] In order to achieve the above-mentioned purpose, the present invention provides the following technical solutions: an end wall cooling structure suitable for high-pressure turbine blades, the cooling structure comprising an annular end wall and an end wall mounting edge for mounting the annular end wall, an impact plate is provided on the outer side of the annular end wall, a plurality of impact holes are arranged on the impact plate, an independent impact cavity is provided in the middle layer of the annular end wall in the leading edge area through a partition, a plurality of spoiler ribs are provided in the impact cavity, a plurality of first air film holes are arranged in the leading edge area of ​​the annular end wall, the end wall mounting edge is located upstream of the annular end wall, and air vents are distributed on the wall surface in contact with the impact plate on the outer side of the end wall, an air collecting cavity for collecting cold air flowing out of the air vents is provided inside the end wall mounting edge, and a plurality of second air film holes and a plurality of third air film holes are opened on the air inlet edge of the inner side of the end wall near the air collecting cavity.

[0006] The end wall cooling structure for high-pressure turbine blades provided by the present invention also has the following characteristics: the impact cavity divides the leading edge high-pressure area into different areas according to the pressure distribution inside the end wall.

[0007] The end wall cooling structure for high-pressure turbine blades provided by the present invention also has the following characteristics: the diameter D of the impact hole is 0.6-0.8 mm; the area of ​​the impact hole is 1.1-1.3 times the area of ​​the first air film hole; the impact hole and the spoiler ribs are arranged in a fan-shaped staggered manner.

[0008] The end wall cooling structure for high pressure turbine blades provided by the present invention also has the following characteristics: a plurality of impact holes are arranged in a fan-shaped manner in five rows, and the position close to the leading edge is set as the first row.

[0009] The lateral spacing between the first and second rows of impact holes is P1 = 2.8*D-3.5*D, and the lateral spacing between the third, fourth and fifth rows of impact holes is P2 = 3.4*D-4.5*D.

[0010] The longitudinal spacing between the first row and the second row is L4 = 2.5*D-4.2*D, the longitudinal spacing between the second row and the third row is L5 = 3.4*D-7.0*D, the longitudinal spacing between the third row and the fourth row is L6 = L5 = 3.4*D-7.0*D, and the longitudinal spacing between the fourth row and the fifth row is L7 = 4.2*D-8.4*D.

[0011] The end wall cooling structure suitable for high-pressure turbine blades provided by the present invention also has the following characteristics: the angle α between the air vent and the wall surface of the end wall mounting edge is not less than 50°, and the area of ​​the air vent is 1.2-1.4 times the sum of the areas of the second air film hole and the third air film hole.

[0012] The end wall cooling structure suitable for high-pressure turbine blades provided by the present invention also has the following characteristics: the cross-section of the air collecting cavity is crescent-shaped, the cross-sectional width L2 is 50%-60% of the end wall mounting edge width L1, and the height L3 of the air collecting cavity is 2*L2-2.5*L2.

[0013] The end wall cooling structure for high-pressure turbine blades provided by the present invention also has the following characteristics: the second film hole and the third film hole are separated by a leading edge stagnation point, the second film hole is inclined toward the blade basin side, and the third film hole is inclined toward the blade back side, and the angle β between the second film hole and the third film hole is 25°-35°.

[0014] The included angle between the axis of the second air film hole and the third air film hole and the inner wall surface of the end wall is 30°-35°.

[0015] The second air film hole and the third air film hole include one or more holes selected from round holes and irregular-shaped holes.

[0016] The end wall cooling structure suitable for high-pressure turbine blades provided by the present invention also has the following characteristics: the angle between the hole axis of the first air film hole and the inner wall surface of the end wall is 30°-35°, and multiple first air film holes are evenly arranged along the leading edge blade profile.

[0017] Beneficial effects

[0018] The novel endwall cooling structure for high-pressure turbine blades proposed by this invention utilizes strategically placed discrete film holes in the inner endwall area most affected by the mainstream horseshoe vortex, while also enhancing the impact heat exchange effect on the outer endwall. This allows for efficient utilization of cooling air and provides effective cooling protection for the endwall near the blade leading edge. For endwalls with shorter intake edges, which are more challenging to cool, the installation of side bleed air ensures sufficient cooling near the blade leading edge, enhancing the reliability of endwall cooling. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] In order to more clearly illustrate the technical solutions of the embodiments of the present disclosure, the following briefly introduces the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present disclosure. For ordinary technicians in this field, other drawings can be obtained based on these drawings without any creative work.

[0020] Figure 1 A schematic structural diagram of a high-pressure turbine blade end wall provided by an embodiment of the present invention;

[0021] Figure 2 A schematic structural diagram of an end wall cooling structure provided by an embodiment of the present invention;

[0022] Figure 3A schematic structural diagram of an end wall mounted edge gas collecting cavity provided by an embodiment of the present invention;

[0023] Figure 4 A schematic diagram of the structure of the impact hole and the spoiler rib on the outer side of the end wall provided by an embodiment of the present invention;

[0024] Figure 5 This is a schematic diagram of the arrangement of the end wall air film holes provided in an embodiment of the present invention.

[0025] Among them, 1-annular end wall, 2-end wall mounting edge, 11-impact plate, 12-impact hole, 13-impact cavity, 14-spoiler rib, 15-first air film hole, 16-inner air inlet edge of end wall, 21-vent, 22-gas collecting cavity, 23-second air film hole, 24-third air film hole. DETAILED DESCRIPTION

[0026] The present invention is further described in detail below with reference to the accompanying drawings and examples. However, it should be noted that these embodiments are not limitations of the present invention, and any equivalent transformations or substitutions in functions, methods, or structures made by ordinary technicians in this field based on these embodiments are within the scope of protection of the present invention.

[0027] In the description of the embodiments of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside" and "outside" and the like to indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the invention.

[0028] Furthermore, the terms "first," "second," "third," etc., are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of the technical features being referred to. Thus, a feature identified with "first," "second," etc., may explicitly or implicitly include one or more of such features. In the description of the present invention, unless otherwise specified, "plurality" means two or more.

[0029] The terms "mounted," "connected," and "connected" should be understood broadly. For example, they can refer to fixed, detachable, or integral connections; mechanical or electrical connections; direct or indirect connections through an intermediary; and internal communication between two components. Those skilled in the art will understand the specific meanings of these terms in the present invention based on specific circumstances.

[0030] like Figure 1-5 As shown, this embodiment provides an end wall cooling structure suitable for high-pressure turbine blades, the cooling structure includes an annular end wall 1 and an end wall mounting edge 2 for mounting the annular end wall 1, an impact plate 11 is provided on the outer side of the annular end wall 1, a plurality of impact holes 12 are arranged on the impact plate 11, an independent impact cavity 13 is provided in the middle layer of the annular end wall 1 in the leading edge area through a partition, a plurality of spoiler ribs 14 are provided in the impact cavity 13, a plurality of first air film holes 15 are arranged in the leading edge area of ​​the annular end wall 1, the end wall mounting edge 2 is located upstream of the annular end wall 1, and air vents 21 are distributed on the wall surface in contact with the impact plate 11 on the outer side of the end wall, an air collecting cavity 22 for collecting cold air flowing out of the air vents 21 is provided inside the end wall mounting edge 2, a plurality of second air film holes 23 and a plurality of third air film holes 24 are opened on the air inlet edge 16 on the inner side of the end wall of the air collecting cavity 22.

[0031] In the above embodiment, a portion of the cooling air from the high-pressure turbine blade enters the interior of the end wall through the impact holes 12 on the outer side of the end wall, impact-cooling the annular end wall 1. After heat exchange with the spoiler ribs 14, it flows out through the discrete first air film holes 15 on the inner side of the annular end wall 1 near the leading edge of the blade, forming an air film covering the surface of the annular end wall 1. The other portion of the cooling air enters the interior of the end wall through the vent holes 21 on the end wall mounting edge 2, flows through the semi-lunar air collecting cavity 22, and is discharged through the second and third air film holes 23 and 24 on the inner side of the end wall. These two portions of cooling air cooperate to effectively protect the blade end wall. This structure, while rationally arranging the discrete air film holes on the inner side of the end wall, enhances the impact heat exchange effect on the outer side of the end wall, providing effective cooling protection for the end wall near the leading edge of the blade. Especially for end walls that are more difficult to cool and have shorter air inlet edges, by installing side air on the end wall, the amount of cooling air on the end wall near the leading edge of the blade is increased, and the air film coverage effect is improved. This can solve the problem of local cooling difficulties on the end wall near the leading edge of the high-pressure turbine blade, thereby greatly improving the temperature bearing capacity of the turbine blade end wall.

[0032] In some embodiments, the impact cavity 13 divides the leading edge high-pressure area into different zones based on the pressure distribution inside the end wall, which can effectively control the distribution of cold air on the end wall and ensure that the leading edge high-temperature area has sufficient cold air.

[0033] In some embodiments, the diameter D of the impact holes 12 is 0.6-0.8 mm; the area of ​​the impact holes 12 is 1.1-1.3 times the area of ​​the first air film holes 15; and the impact holes 12 and the spoiler ribs 14 are arranged in a staggered fan-shaped pattern. The area of ​​the impact holes 12 refers to the sum of the areas of all impact holes 12, and the area of ​​the first air film holes 15 refers to the sum of the areas of all first air film holes 15.

[0034] In some embodiments, the plurality of impingement holes 12 are arranged in a fan-shaped pattern in five rows, with the first row being the one closest to the leading edge. The lateral spacing P1 between the first and second rows of impingement holes is 2.8*D-3.5*D, the lateral spacing P2 between the third, fourth, and fifth rows of impingement holes is 3.4*D-4.5*D, the longitudinal spacing L4 between the first and second rows is 2.5*D-4.2*D, the longitudinal spacing L5 between the second and third rows is 3.4*D-7.0*D, the longitudinal spacing L6 between the third and fourth rows is 3.4*D-7.0*D, and the longitudinal spacing L7 between the fourth and fifth rows is 4.2*D-8.4*D. This arrangement of impingement holes 12 is used to control cold air, effectively cooling the leading edge stagnation zone.

[0035] In some embodiments, the angle α between the vent holes 21 and the end wall mounting edge 2 is no less than 50°. The area of ​​the vent holes 21 is 1.2-1.4 times the combined area of ​​the second and third air film holes. The maximum angle α is determined based on the specific structure of the air outside the impact plate. The area occupied by all vent holes 21 is 1.2-1.4 times the combined area of ​​all second and third air film holes 23, 24. The combined area of ​​the vent holes 21 ensures that cold air can effectively flow out of the second and third air film holes 23, 24.

[0036] In some embodiments, Figure 1 As shown, the plenum chamber 22 has a crescent-shaped cross-section (AA), with a cross-sectional width L2 that is 50%-60% of the width L1 of the end wall mounting edge 2. The height L3 of the plenum chamber 22 is equal to 2*L2-2.5*L2. This crescent-shaped plenum chamber 22 can collect cold air while reducing vortexes within the chamber, thereby minimizing pressure loss.

[0037] In some embodiments, the second air film hole 23 and the third air film hole 24 are divided by the leading edge stationary point, the second air film hole 23 is inclined toward the blade basin side, and the third air film hole 24 is inclined toward the blade back side, and the angle β between the second air film hole 23 and the third air film hole 24 is 25°-35°.

[0038] The included angle between the axis of the second air film hole 23 and the third air film hole 24 and the inner wall surface of the end wall is 30°-35°.

[0039] The second air film hole 23 and the third air film hole 24 include one or more holes selected from round holes and irregular-shaped holes.

[0040] In some embodiments, the included angle between the axis of the first film holes 15 and the inner surface of the end wall is 30°-35°, and multiple first film holes 15 are evenly distributed along the leading edge airfoil. The interaction between the first film holes 15, the second film holes 23, and the third film holes 24 improves the air film coverage of the leading edge area.

[0041] In some embodiments,

[0042] The division of the independent impact cavity 13 outside the annular end wall 1 is based on the pressure distribution inside the end wall, and the leading edge high pressure area is independently divided.

[0043] The spoiler ribs 14 and the impact holes 12 on the outer side of the annular end wall 1 are arranged alternately in a fan shape.

[0044] The arrangement of the impact holes 12 on the outer side of the annular end wall 1 is such that the lateral spacing P1 of the two rows near the leading edge is 1.9 mm, and the lateral spacing P2 at other positions is 2.4 mm; the longitudinal spacing of each row is L4 = 2.0 mm, L5 = 3.0 mm, L6 = 3.0 mm, and L7 = 3.8 mm respectively.

[0045] The diameter of the impact hole 12 on the outer side of the annular end wall 1 near the leading edge of the blade is 0.6 mm, and the area of ​​the impact hole 12 is 1.1 times the area of ​​the first air film hole 15 on the inner side of the end wall 1 .

[0046] The included angle α between the vent hole 21 and the wall surface of the end wall mounting edge 2 is 50°, and the area of ​​the vent hole 21 is 1.1 times the sum of the areas of the second air film hole 23 and the third air film hole 24 .

[0047] The cross-sectional width L2 of the gas collecting cavity 22 along the AA direction accounts for 50% of the width L1 of the end wall mounting edge 2 , and the height L3 of the gas collecting cavity 22 is equal to L2*2.

[0048] The included angle between the axis of the first air film hole 15 , the second air film hole 23 and the third air film hole 24 on the inner side of the annular end wall 1 and the inner wall surface of the end wall 1 is 30°.

[0049] The air film holes 23 and 24 on the inner side of the annular end wall 1 are circular holes, with the leading edge stagnation point as the dividing point, and are inclined toward the blade basin and the blade back side respectively, with an angle β of 30°.

[0050] In some embodiments,

[0051] The division of the independent impact cavity 13 outside the annular end wall 1 is based on the pressure distribution inside the end wall, and the leading edge high pressure area is independently divided.

[0052] The spoiler ribs 14 and the impact holes 12 on the outer side of the annular end wall 1 are arranged alternately in a fan shape.

[0053] The arrangement of the impact holes 12 on the outer side of the annular end wall 1 is such that the lateral spacing P1 of the two rows near the leading edge is 2.0 mm, and the lateral spacing P2 at other positions is 2.4 mm; the longitudinal spacings of each row are L4 = 2.2 mm, L5 = 3.0 mm, L6 = 3.0 mm, and L7 = 4.0 mm, respectively.

[0054] The diameter of the impact hole 12 on the outer side of the annular end wall 1 near the leading edge of the blade is 0.8 mm, and the area of ​​the impact hole 12 is 1.3 times the area of ​​the first air film hole 15 on the inner side of the annular end wall 1 .

[0055] The included angle α between the vent hole 21 and the wall surface of the end wall mounting edge 2 is 60°, and the area of ​​the vent hole 21 is 1.3 times the sum of the areas of the second air film hole 23 and the third air film hole 24 .

[0056] The cross-sectional width L2 of the gas collecting cavity 22 along the AA direction accounts for 60% of the width L1 of the end wall mounting edge 2 , and the height L3 of the gas collecting cavity 22 is L2*2.2.

[0057] The included angle between the axis of the first air film hole 15 , the second air film hole 23 and the third air film hole 24 on the inner side of the annular end wall 1 and the inner wall surface of the end wall 1 is 35°.

[0058] The second air film hole 23 and the third air film hole 24 on the inner side of the annular end wall 1 are special-shaped holes, with the leading edge stagnation point as the dividing point, and are inclined toward the blade basin and the blade back side respectively, with an angle β of 30°.

[0059] The above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention shall be included within the scope of protection of the present invention. The above description is merely a preferred embodiment of the present invention. It should be noted that those skilled in the art can make various improvements and variations without departing from the technical principles of the present invention, and such improvements and variations shall also be considered within the scope of protection of the present invention.

Claims

1. An end wall cooling structure suitable for high pressure turbine blades, characterized in that: The cooling structure includes an annular end wall and an end wall mounting edge for mounting the annular end wall. An impact plate is provided on the outer side of the annular end wall, and a plurality of impact holes are arranged on the impact plate. An independent impact cavity is provided in the leading edge area of ​​the middle layer of the annular end wall through a partition. A plurality of spoiler ribs are provided in the impact cavity. A plurality of first air film holes are arranged in the leading edge area of ​​the annular end wall. The end wall mounting edge is located upstream of the annular end wall, and the wall surface in contact with the impact plate outside the end wall is provided with vent holes. An air collecting cavity for collecting the cold air flowing out of the vent holes is provided inside the end wall mounting edge. The gas collecting cavity is provided with a plurality of second gas film holes and a plurality of third gas film holes on the gas inlet edge close to the inner side of the end wall. The impact cavity divides the leading edge high pressure area into different areas according to the pressure distribution inside the end wall. Part of the cold air from the high-pressure turbine blades enters the interior of the end wall through the impact holes on the outside of the end wall, impact-cools the annular end wall, and after heat exchange through the spoiler ribs, flows out through the discrete first air film holes on the inside of the annular end wall near the leading edge of the blade, forming an air film covering the surface of the annular end wall; the other part of the cold air enters the interior of the end wall from the air vents on the mounting edge of the end wall, flows through the semi-lunar air collecting cavity, and is discharged from the second air film holes and the third air film holes on the inside of the end wall. The two parts of cold air cooperate with each other to achieve effective protection of the blade end wall.

2. The end wall cooling structure for high pressure turbine blades according to claim 1, characterized in that: The diameter D of the impact hole is 0.6-0.8 mm; the area of ​​the impact hole is 1.1-1.3 times the area of ​​the first air film hole; the impact hole and the spoiler ribs are arranged in a fan-shaped staggered manner.

3. The end wall cooling structure for high-pressure turbine blades according to claim 2, characterized in that: There are five rows of impact holes arranged in a fan shape, and the position close to the leading edge is set as the first row. The horizontal spacing between the first and second rows of impact holes is P1 = 2.8*D-3.5*D, and the horizontal spacing between the third, fourth and fifth rows of impact holes is P2 = 3.4*D-4.5*D. The longitudinal spacing between the first row and the second row is L4=2.5*D-4.2*D, the longitudinal spacing between the second row and the third row is L5=3.4*D-7.0*D, the longitudinal spacing between the third row and the fourth row is L6=L5=3.4*D-7.0*D, and the longitudinal spacing between the fourth row and the fifth row is L7=4.2*D-8.4*D.

4. The end wall cooling structure for high pressure turbine blades according to claim 1, characterized in that: The included angle α between the vent hole and the wall surface of the end wall mounting edge is not less than 50°, and the area of ​​the vent hole is 1.2-1.4 times the sum of the areas of the second air film hole and the third air film hole.

5. The end wall cooling structure for high pressure turbine blades according to claim 1, characterized in that: The cross section of the gas collecting cavity is half-moon shaped, the cross section width L2 is 50%-60% of the width L1 of the end wall mounting edge, and the height L3 of the gas collecting cavity is 2*L2-2.5*L2.

6. The end wall cooling structure for high pressure turbine blades according to claim 1, characterized in that: The second air film hole and the third air film hole are divided by the leading edge stationary point. The second air film hole is inclined toward the leaf basin side, and the third air film hole is inclined toward the leaf back side. The angle β between the second air film hole and the third air film hole is 25°-35°. The included angle between the axis of the second air film hole and the third air film hole and the inner wall surface of the end wall is 30°-35°. The second air film hole and the third air film hole include one or more holes selected from round holes and irregular-shaped holes.

7. The end wall cooling structure for high pressure turbine blades according to claim 1, characterized in that: The included angle between the hole axis of the first air film hole and the inner wall surface of the end wall is 30°-35°, and the plurality of first air film holes are evenly arranged along the leading edge blade profile.

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